Method and a system for putting a space vehicle into orbit, using thrusters of high specific impulse

ABSTRACT

The method serves to place a space vehicle, such as a satellite, on a target orbit such as the orbit adapted to normal operation of the space vehicle and starting from an elliptical initial orbit that is significantly different from, and in particular more eccentric than the target orbit. The space vehicle is caused to describe a spiral trajectory made up of a plurality of intermediate orbits while a set of high specific impulse thrusters mounted on the space vehicle are fired continuously and without interruption, thereby causing the spiral trajectory to vary so that on each successive revolution, at least during a first stage of the maneuver, perigee altitude increases, apogee altitude varies in a desired direction, and any difference in inclination between the intermediate orbit and the target orbit is decreased, after which, at least during a second stage of the maneuver, changes in perigee altitude and in apogee altitude are controlled individually in predetermined constant directions, while any difference in inclination between the intermediate orbit and the target orbit continues to be reduced until the apogee altitude, the perigee altitude, and the orbital inclination of an intermediate orbit of the space vehicle have substantially the values of the target orbit.

This is a divisional application of prior application Ser. No.08/833,094 filed Apr. 4, 1997, entitled: A METHOD AND A SYSTEM FORPUTTING A SPACE VEHICLE INTO ORBIT, USING THRUSTERS OF HIGH SPECIFICIMPULSE now U.S. Pat. No. 6,116,543.

FIELD OF THE INVENTION

The present invention relates to a method of putting a space vehicle,such as a satellite, into a target orbit such as the orbit for normaloperation of the space vehicle, starting from an elliptical initialorbit that is significantly different and in particular more eccentricthan the target orbit.

PRIOR ART

Most artificial satellites are now fitted with a thruster systemenabling them to move in space, in particular to correct imperfectionsof trajectory, due in particular to imperfections with which they wereput into orbit, to the gravitational attraction of the moon and the sun,to potential effects due to the earth being non-spherical, toaerodynamic, magnetic, and electrical effects associated with the earth,and to the action of solar radiation. The thrust system of a satellitealso enables the satellite to be put on station, to make changes to itsorbit, to ensure that it is appropriately oriented in space, or indeedto ensure that the attitude control system remains functional byenabling the inertia wheels fitted to such satellites to be desaturated.

Such thrust systems generally give a satellite the ability to move inany direction, with one direction being generally preferred forperforming movements of large amplitude.

In terms of mass budget, the satellite's thruster system generallyconstitutes a major component, or even the largest component.

Most of the thruster industry became interested very early on intechniques for reducing the mass of a thruster system. Given thatspecific impulse is a characteristic value of a thruster, specifying theimpulse provided per unit mass ejected or consumed, high specificimpulse thrusters have been designed, developed, and evaluated. By wayof example, mention can be made of “resistojet” type thrusters, closedelectron drift plasma thrusters, FEEP field emission thrusters, ionbombardment thrusters, and heliothermal thrusters.

Theocratically speaking, increasing specific impulse is based ontransforming zero-mass power, i.e. power whose production consumespractically no matter, into mechanical power applied to particles ofmatter. In practice, such zero-mass power is electrical power or thermalpower obtained from solar radiation, or indeed as obtained from aradioisotopic generator.

The thrust obtained from high specific impulse thrusters of this kindtherefore depends on the level of electrical or thermal power that canbe supplied to them. On a satellite, such power is limited by the sizeof its solar panels, by the size of its solar thermal energyconcentrators, or by the size of its radioisotopic generator, or indeedby the size of its energy storage means. As a result, the thrustdelivered by any high specific impulse thruster is small or very smallcompared with the thrust from a conventional chemical engine, e.g. 400 N(a typical value for a satellite apogee engine).

The greater the electrical or thermal power transformed into mechanicalpower applied to a given mass of particles, the greater the resultingspecific impulse. Thus, the greater the specific impulse of thrusters,the lower the thrust provided for given consumption of electrical orthermal power. This property is substantially valid for all types ofhigh specific impulse thruster.

This property has the following effects on propulsion systems: for agiven total delivered impulse (i.e. the cumulative value or the timeintegral of the force delivered to the vehicle over the entire durationof firing), there is both a remarkable reduction in the mass of matterconsumed by higher specific impulse thrusters and a correspondingincrease in the time over which such thrusters operate.

Thrust of the high specific impulse type is suitable for the maneuversperformed when a satellite is on its nominal operating orbit, since theforces that need to be delivered are then low or very low, thus makingit possible to achieve real advantages over chemical thrust systems (oflower specific impulse).

The opposite applies when a satellite is initially placed on an orbitthat is very different from its nominal orbit and the satellite needs tooperate its own thrust system to move from its initial orbit to itsnominal orbit.

Under such circumstances, the total duration of the transfer maneuvertends to be lengthy whereas, on the contrary, it would be better to beable to minimize said duration. The longer the total duration of thetransfer maneuver, the greater the financial burden, and costsassociated with putting the vehicle into orbit also increase (includingground station costs and the cost of tracking teams on the ground).Also, a long duration transfer maneuver increases the risk of the spacevehicle being damaged as it passes through the Van Allen belts (whichvary in position, but which may be situated, for example, in thevicinity of the following altitudes: 1800 km, 2000 km, 10,000 km, and21,000 km).

It is desirable to minimize the number of revolutions the vehicleperforms on orbits that pass through the Van Allen belts, in particularto minimize the additional hardening of components or solar cells thatwould otherwise be needed to protect them against the electromagneticwaves or radiation emitted by the protons or electrons present in thebelts.

Various examples of satellite maneuvers have already been proposed thatmake use of thrusters having high specific impulse and low thrust.

Thus, the article by A. G. Schwer, U. W. Schöttle, E. Messerschmid ofthe University of Stuttgart (Germany), published at the 46thInternational Astronautic Congress in 1995 and entitled “Operationalimpacts and environmental effects on low-thrust transfer missions oftelecommunication satellites”, shows that a transfer maneuver between aconventional initial orbit constituted by the geostationary transferorbit (GTO) of the Ariane 4 rocket and a final orbit constituted by ageostationary orbit (GEO). The transfer maneuver comprises a largenumber of thrust arcs about the apogee of the GTO orbit created by“Arcjet” type thrusters having high specific impulse, such that theorbit of the satellite deforms progressively until it reaches the finalgeostationary orbit GEO.

Document EP-B-0 047 211 (inventor A. Mortelette) also describes a methodof changing orbit by thrust arcs.

For the two methods described in the two above-mentioned documents,apogee altitude is constrained to remain constant or to vary veryslowly. The duration required by the maneuvers described is quite large,such that to reduce said duration it is necessary to increase thrust.Also, the number of times the engine needs to be started is also largeand this can give rise to major operational constraints. When thesatellite is on an orbit with a variable sidereal period which ischanging constantly as a function of the extent to which its position isprogressing, and which can be different or even very different from thesidereal period of earth rotation, the satellite is not always visiblefrom a given ground station when it is necessary to start its thrusters.This means that any procedure for putting the satellite into orbit thatrequires thrusters to be started on numerous occasions cannot beperformed safely using only one ground station. On the contrary, it mustbe possible to make use of a plurality of ground stations located atdifferent places throughout the time required for putting the satelliteinto orbit. The cost of ground operations and rental of ground stationsis not negligible.

In order to reduce the total duration of the maneuver in the schemes putforward in the above-mentioned document for putting a satellite intoorbit, it would be better to have high thrust thrusters. However, undersuch circumstances, for given power delivered to the thrusters, thespecific impulse would have to be smaller and consequently the massconsumed during the maneuver would be greater. The various solutionsthat have already been recommended for putting a satellite into orbit inthis manner are therefore of relatively low performance.

Proposals have already been made by Irving, to pass from an initialcircular orbit to a final circular orbit by continuously operatingthrusters that are in alignment either with the local horizontal, orelse with the orbital speed (speed relative to earth) so that the orbitdeforms progressively and comes to the immediate vicinity of the finaltarget orbit, or even reaches it. That type of maneuver leads to aspiral type trajectory and the maneuver can be performed by starting thethruster once only. Nevertheless, the number of revolutions in the VanAllen belts cannot be optimized, which is a drawback, and above all, atpresent that type of maneuver applies only to transfers between orbitsthat are circular or, where appropriate, between particular ellipticalorbits.

Proposals have also been made to combine a set of maneuvers by means ofthrust arcs with a subsequent maneuver of the spiral type, in thecontext of a particular maneuver described in document EP-A-0 673 833(invention of A. Spitzer). If use is made only of high specific impulseand low thrust propulsion, that technique for putting a satellite intoorbit is of relatively low performance, particularly in terms ofduration. The number of revolutions that pass through the Van Allenbelts is high and the number of times the thrusters are started islarge, and unfortunately this large number of thruster starts isnecessary during the first stage during which perigee altitude isincreased, and in which the sidereal period is, in particular, differentfrom the sidereal period of rotation of the earth.

A. Spitzer has also proposed a solution using “hybrid” propulsion, i.e.combining conventional chemical engines operated in a first maneuveringstage over a plurality of thrust arcs, and high specific impulsethrusters implemented in a second maneuver stage that is spiral shaped.Under such circumstances, the total duration of the transfer maneuver isgreatly reduced and the number of revolutions in the Van Allen belts issmaller. However, the mass consumed during the maneuver is relativelylarge, and above all two different types of propulsion system need to bepresent, thereby increasing cost compared with systems having only onetype of propulsion system, making the architecture of the satellite morecomplex, and increasing preparation and launch costs, in particularbecause of the need to fill various tanks with different substances andbecause of the precautions that need to be taken against the risks ofpollution or of fire.

OBJECT AND BRIEF DESCRIPTION OF THE INVENTION

The present invention seeks to remedy the above-mentioned drawbacks and,in particular, to enable a satellite placed by a launcher or a spacevehicle on an orbit that is not the nominal operational orbit of thesatellite, and to reach said orbit in a manner that is particularlyeffective while using thrusters of high specific impulse and low thrust.

The invention seeks in particular to make it possible to minimize theduration of the transfer maneuver from the initial orbit to a targetorbit.

The invention also seeks to improve the reliability of the maneuvers.

Another object of the invention consists in reducing the cost ofmanufacturing the space vehicle and of using it, and of reducing therisks associated with the use of chemical substances.

These objects are achieved by a method of placing a space vehicle, suchas a satellite, on a target orbit such as that adapted to normaloperation of the space vehicle and starting from an elliptical initialorbit that is substantially different from the target orbit and inparticular more eccentric than the target orbit, wherein the spacevehicle is caused to describe a spiral trajectory made up of a pluralityof intermediate orbits during single continuous firing of a set of highspecific impulse thrusters mounted on the space vehicle, progress of thespiral trajectory being controlled in such a manner that on eachsuccessive revolution, at least during a first stage of the maneuver,perigee altitude increases, apogee altitude moves in a determineddirection, and any difference in inclination between an intermediateorbit and the target orbit decreases, then, at least during a secondstage of the maneuver, changes in perigee and apogee altitude arecontrolled individually in constant predetermined directions while anydifference in inclination of an intermediate orbit relative to thetarget orbit continues to be reduced until the apogee altitude, theperigee altitude, and the orbital inclinations of an intermediate orbitof the space vehicle have reached substantially the values of the targetorbit.

Where appropriate, the assembly of high specific impulse thrusters maycomprise a single high specific impulse thruster.

In this way, only one type of thruster is required for putting thesatellite into its target orbit.

However, it is not impossible to include auxiliary thrusters of othertypes, such as cold gas thrusters or resistojets, and making use of thesame gas as the high specific impulse thrusters, e.g. xenon, therebymaking it possible during short instants to obtain greater thrustlevels, while avoiding the above-mentioned drawbacks of chemical thrust.

Such auxiliary thrusters may be used in an initial stage, e.g. forcontrolling parasitic moments due to deploying the solar panels.

In a first particular implementation, for a space vehicle initially onan elliptical orbit that is substantially different from the targetfinal orbit for normal operation of the vehicle, during the first stageof the maneuver starting from the beginning of continuous firing of thethrusters, and on each successive revolution, apogee altitude ofintermediate orbits is caused to increase and perigee altitude toincrease to a lesser extent, and during the second stage of themaneuver, until the end of continuous firing, on each successiverevolution, apogee altitude of intermediate orbits is decreased whileperigee altitude is increased.

In a second particular implementation-, during the first stage of themaneuver, from the beginning of continuous firing of the thrusters, andon each successive revolution, apogee altitude of the intermediateorbits is increased and so is perigee altitude, during a second stage ofthe maneuver, near the middle of continuous firing, and on eachsuccessive revolution, apogee altitude of intermediate orbits isdecreased and perigee altitude is increased, and then during a thirdstage of the maneuver, once the eccentricity of intermediate orbits ofthe space vehicle has substantially reached that of the target orbit,and until the end of continuous firing, during each successiverevolution, apogee altitude is decreased and erigee altitude isdecreased while any difference of inclination of intermediate orbitsrelative to the target orbit continues to be decreased, until apogeealtitude, perigee altitude, and orbital inclination of an intermediateorbit of the space vehicle have reached substantially the values of thetarget orbit.

In a third particular implementation, during the first stage of themaneuver starting from the beginning of continuous firing of thethrusters, and on each successive revolution, apogee altitude of theintermediate orbits is decreased and perigee altitude is increased, andthen during the second stage of the maneuver, until the end ofcontinuous firing, once the eccentricity of the intermediate orbits ofthe space vehicle has reached substantially that of the target orbit,and on each successive revolution, apogee altitude of the intermediateorbits is decreased and perigee altitude is also decreased.

The invention also provides a system for placing a space vehicle, suchas a satellite, on a target orbit such as that adapted to normaloperation of the space vehicle and starting from an elliptical initialorbit that differs substantially from the target orbit, and inparticular that has eccentricity that is different from that of thetarget orbit, the system comprising:

a set of platforms mounted on the space vehicle;

a set of thrusters mounted on said platforms and

having high specific impulse, greater than 5000 Ns/kg, and low thrust,less than 10 N, to create a total thrust force applied to the spacevehicle;

a control device for putting the thrusters into continuous operationafter the space vehicle has been placed on its initial orbit so as toenable said space vehicle to reach a target orbit via a completelyspiral trajectory, ignoring possible service interruptions, and to stopthe thrusters firing when the apogee altitude, the perigee altitude, andthe orbital inclination of an intermediate orbit of the space vehicleare substantially equal to the values of the target orbit; and

a thrust direction control device comprising at least first means foroperating, during a first stage of continuous firing of the set ofthrusters, to generate first control signals for aiming total thrustsuch that on each successive revolution of the space vehicle and in eachintermediate orbit, apogee altitude increases, perigee altitudeincreases to a lesser extent, and any difference in the inclination ofthe intermediate orbit relative to the target orbit is decreased, and atleast second means for operating, during a second stage of the set ofthrusters firing, to generate second control signals for aiming totalthrust such that on each successive revolution of the space vehicle andin each intermediate orbit, apogee altitude decreases, perigee altitudeincreases, and any difference of inclination between the intermediateorbits and the target orbit decreases.

Advantageously, the set of thrusters having high specific impulsecomprises plasma type closed electron drift thrusters, ion thrusters, or“Arcjet” type thrusters (i.e. electric arc thrusters).

In a particular embodiment, said first and second means for generatingthe first and second control signals for aiming total thrust include aset of sensors enabling the attitude of the space vehicle to be verifiedor checked, and are actuated while the device for controlling operationof the thrusters in continuous manner continues to act so that firing iscontinuous.

More particularly, said first means of said device for controllingthrust aiming include means for aligning total thrust in a localhorizontal plane, mainly directed in the direction of the speed of thespace vehicle.

In which case, more particularly, said second means of said device forcontrolling thrust aiming comprise means for operating around apogee tobring the total thrust into alignment in a local horizontal plane mainlywith the speed direction of the space vehicle, and means for operatingaround perigee to bring the total thrust into alignment with a directionopposite to the orbital speed of the space vehicle in a plane that issubstantially orthogonal to the orbital plane.

The means for operating around apogee to align total thrust serve to aimtotal thrust substantially along the osculating half-ellipse centered onapogee.

The means for operating around perigee to align total thrust serve toaim total thrust substantially on the osculating half-ellipse centeredon perigee.

In a variant embodiment, said first and second means of said device forcontrolling thrust aiming include means for aligning thrust in a planepassing via directions that are fixed or almost fixed in space andpassing almost perpendicularly to the plane that is tangential to theorbit.

The system may include means for aiming total thrust that areconstituted by means belonging to the space vehicle such as inertiawheels of a system for controlling the attitude of the space vehicle.

The system may also include means for aiming total thrust constituted byat least some of said platforms for supporting sets of thrusters, saidplatforms being steerable. Adjustable-thrust thrusters can also be used.

The system of the invention may include means for steering total thrustincluding means for differentially aiming said steerable platforms andmeans for servo-controlling the thrust from each thruster to apredetermined value, thereby enabling total thrust to be generated thatpasses through the center of mass of the space vehicle and that has acomponent lying outside the orbital plane.

Advantageously, the steerable platforms can be teered through more than10° at least about one axis.

In an advantageous embodiment of the invention, said thrusters havinghigh specific impulse also constitute means for controlling attitude andorbit of the space vehicle such as a satellite.

It may be observed that with the method of the invention, by ensuringboth a continuous increase of perigee (possibly terminating by adecrease of altitude), and a variation in apogee altitude beginning withan increase and ending with a decrease, it is possible from almost anyelliptical initial orbit to reach a very diverse multitude of finalorbits, while keeping the thrusters continuously in operation.

In particular, when the initial orbit and the final target orbit haveapogees that are close together, it is possible according to theinvention to ensure that any increase in apogee altitude during thrusterfiring is substantially equal to any decrease. In which case, the entiremaneuver ends up with an increase in perigee altitude.

Although the type of maneuver defined in the method and system of theinvention can be applied to elliptical initial orbits that are arbitraryrelative to the target orbit, the present invention is particularlyadvantageous when the initial orbit is elliptical, fairly eccentric(having eccentricity greater than 0.2), of sidereal period that isshorter than that of the target orbit (which implies in particular thatthe perigee altitude of the initial orbit is lower than that of thetarget final orbit), that has its apogee close to that of the targetorbit, and in which the target final orbit is circular or has noeccentricity (eccentricity less than 0.1). Such circumstances apply whenlaunching a geostationary satellite into a geostationary transfer orbit,or when launching into its transfer orbit a satellite that is to occupya circular orbit at medium altitude (e.g. 20,000 km). Successiveincreases of apogee during firing lead to best efficiency for themaneuver which consists mainly in increasing perigee altitude. Towardsthe end of thruster firing, successive decreases in apogee altitude leadto additional consumption by the thrusters, but such additionalconsumption remains modest and enables a significant reduction to beachieved overall concerning the total duration of the maneuver.

The method of the invention can be applied advantageously when theinitial orbit is elliptical and has a sidereal period that is smallerthan that of the target final orbit, even when the apogee altitude ofthe initial orbit is very different from that of the target orbit andthe target final orbit is circular.

The third implementation of the invention is particularly adapted tominimizing the number of revolutions in the Van Allen belts. In thismethod, the high specific impulse thrusters are fired continuously andtheir total thrust is steered, e.g. to increase perigee altitude and todecrease apogee altitude of the intermediate orbits, and then withoutinterrupting thruster operation, their total thrust is steered so as tocontinuously decrease apogee altitude and simultaneously decreaseperigee altitude such that, in the end, perigee and apogee altitudescoincide with those of the target final orbit.

Such a particular implementation is well suited to launching a spacevehicle onto a transfer orbit that is very highly eccentric and of asidereal period greater than or equal to that of the target orbit. Anextreme characteristic example is launching a satellite that is tooccupy a geostationary orbit having a period of 1 day, starting from anorbit with a 620 kg perigee and a 330,000 km apogee, i.e. an orbithaving an 8-day period. With an initial mass of 2950 kg and a totalthrust of 0.64 N, it is possible to operate the high specific impulsethrusters of the invention so as to have only two revolutions in the VanAllen belts. By using such a small number of intermediate orbits passingthrough the Van Allen belts, the performance of the method of theinvention turns out in this case to be as good as that of conventionalmethods implementing high thrust propulsion.

One of the advantages of the method and the system of the invention isthat the maneuver serving, for example, to move a satellite from aconventional geostationary transfer orbit (GTO) to geostationary orbit(GEO) while starting the thrusters only once, constitutes an importantadvantage since starting operations are always relatively difficult. Itis thus desirable to limit starting operations given that starting ischaracteristic of a transient state and that always requires a greatdeal of attention. For example, with an electrical thruster, it takesseveral minutes to prepare the thruster and put it into operation, inapplication of a particular sequence of events. Also, when a pluralityof thrusters are used to perform the maneuver, it is necessary to beable to ensure that the various thrusters all start simultaneously,particularly if their individual thrusts do not pass close to the centerof mass of the space vehicle. If they do not start simultaneously, thenit is necessary to act on the attitude control system of the spacevehicle to cancel attitude drift. The control system can becomesaturated and starting must be aborted and tried again later. It is thusparticularly advantageous, throughout the duration of travel between theinitial orbit and the target orbit, and in particular at least while thesidereal periods of the intermediate orbits are significantly differentfrom the sidereal period of rotation of the earth, to be able to ensurethat the thrusters need starting only once. Such a single start can beperformed when the satellite is well placed for visibility from aparticular ground station. There is thus no need to have a plurality ofground stations available, and in addition, the location at which thesingle start is performed is of relatively little importance. Further,this advantage of starting only once shows that the method of theinvention for putting a satellite into its target orbit is well suitedto automating control, and consequently to the space vehicle beingautonomous while it is being put into orbit, thereby reducing the costsof putting it into orbit.

Another important advantage of the present invention lies in the factthat a complete maneuver for putting the satellite into orbit makes useof only one type of thruster, i.e. thrusters of high specific impulseor, where appropriate, also of auxiliary thrusters but that make use ofthe same inert gas as said thrusters of high specific impulse. Thiscontributes to reducing the cost of making the space vehicle and ofoperating it, and it limits the risks associated with using chemicalsand hypergols or other toxic substances, given that the main types ofhigh specific impulse thruster hardly ever use chemicals other than aninert gas such as xenon.

Also, since the maneuver of the invention necessarily leads to thethrusters operating continuously, the mass consumed by the thrustersduring the maneuver is directly proportional to the duration of themaneuver. As a result, in the context of the present invention, itsuffices to optimize the steering of the thrusters during the maneuverin order to minimize the duration of the maneuver and thus also minimizethe mass that is consumed during the maneuver.

Various optimization methods can be used to determine appropriatesteering relationships while satisfying the proposed criteria forvariations in apogee and perigee parameters.

Thus, in particular, for steering at the beginning of the maneuver, itturns out that a thrust relationship lying in the local horizontal planeis particularly effective from a conventional geostationary transferorbit GTO. Towards the end of the maneuver, a thrust relationship lyingin the horizontal plane around the apogee of the orbit andanti-tangential to the orbital speed around perigee is satisfactory. Afixed steering or inertial relationship is also of interest during themaneuver.

At the end of the maneuver, it may be advantageous to have a satellitefacing towards the earth (consequence of a steering relationship forthrust in the local horizontal plane) in particular for the purpose ofperforming certain adjustments that may be necessary to the payload, forexample, simultaneously with finalization of the maneuver for puttingthe satellite into its orbit (i.e. while the thrusters are stilloperating).

It may be observed that the measures recommended by the presentinvention for steering the thrusters turn out, in the case of reaching ageostationary orbit from a conventional transfer orbit, to beparticularly suited to performing short duration maneuvers withoutsignificantly increasing the mass of matter consumed during saidduration as compared with conventional systems using chemicalpropulsion.

Thus, by way of example, for a 2950 kg satellite fitted with plasmathrusters having a total thrust of 0.64 N, high specific impulse of16,000 Ns/kg, and a specific power of 16 kW/N, with the satellite beinginitially placed on a geostationary transfer orbit of the type intendedfor Ariane 5 launchers (apogee 36,000 km and perigee 620 km), themaneuver lasts for no more than 3.7 months for an electrical powerconsumption of 10 kW and for a mass consumption to transfer the orbit togeostationary orbit of 380 kg.

In the method of the invention for putting a satellite into orbit, it ispossible to consider the notion of the specific impulse of the maneuveras a whole, defined as the ratio of the total impulse delivered toperform the maneuver of putting the satellite into its orbit divided bythe initial mass of the vehicle (Mi in kg), as follows:

I_(spmaneuver)=F×Δt/Mi

Since the thrusters are firing continuously, the total impulse deliveredis equal to the product of the force (in Newtons) multiplied by theduration of the maneuver Δt (in seconds).

In the example under consideration, I_(spmaneuver) is 2070 N.s/kg. Thisquantity depends relatively little on the specific impulse of thethrusters (providing it is high). This notion of specific impulse of themaneuver makes it possible to evaluate rapidly the main characteristicsand advantages of the method of the invention for putting a satelliteinto orbit.

In comparison, a conventional satellite having the same payload andusing conventional chemical propulsion would need to have a launch massof 4100 kg, i.e. one extra (metric) tonne, thus increasing launch costs,or for constant launch costs, reducing payload and thus profitability ofthe space vehicle.

Also, with a satellite having the same payload and using a high specificimpulse type of propulsion giving 16,000 Ns/kg and thrust analogous tothat mentioned in the above example (0.64 N), but achieving orbittransfer by making use, for example, of a method relying on thrust arcsaround orbit apogee or even a method constituted by thrust arcs aroundapogee and thrust arcs around perigee, the takeoff mass would be of thesame order of magnitude (e.g. 2860 kg), but the total duration of themaneuver for putting the satellite on its final orbit would belengthened to 4.8 months, i.e. it would be about 30% longer than thatrequired for putting a satellite into orbit by means of the presentinvention.

The advantage provided in this way by the present invention is verysignificant since it serves in particular to reduce the financial costsdue to the investment that must be made to build the satellite and tolaunch it before it starts earning. In addition, ground-based means areused for less time during the maneuver so the total cost thereof is alsoreduced.

Another advantage provided by the type of maneuver performed in theinvention lies in a significant reduction of the time spent passingthrough the Van Allen belts in comparison with a maneuver relying onthrust arcs. Because the apogee altitude of the intermediate orbits isincreased, it is possible in particular to increase speed in thevicinity of perigee, specifically where the belts of protons andelectrons are to be found, thereby reducing the time spent passingthrough the Van Allen belts compared with performing the transfer usingthe same thrusters but by means of apogee thrust arcs. In addition,perigee altitude increases much more quickly, so the total number ofpasses through the Van Allen belts is also considerably reduced.

Another advantage associated with the present invention consists in itsefficiency increasing with increasing inclination of the initial orbitrelative to the target final orbit. The increase in apogee altitudemakes it possible to correct a difference of inclination moreeffectively, which is a characteristic advantage specific tosupergeosynchronous orbits, for example, and the greater the altitude,the greater the effectiveness. In this case, it is merely necessary alsoto verify that the apogees of the intermediate orbits are indeedsituated in the plane of the target final orbit and that the totalthrust is steered about the apogee so as to reduce difference ininclination, with thrust lying partially outside the orbital plane.Maximum effectiveness of the maneuver is obtained when the steering oftotal thrust does not change any orbital parameter other than apogeealtitude, perigee altitude, and inclination. It follows that when thesatellite is placed on its initial orbit, if the apogee of the initialorbit lies in the plane of the target final orbit, then thatconfiguration will continue, and correcting inclination difference isperformed more effectively. It is possible to take account of theso-called “disturbing” influences of the moon, the sun, and thenon-spherical gravitational potential of the earth to improve theeffectiveness of the correction.

Another advantage associated with the present invention lies in thepossibility of deploying the satellite completely or in part before itsthrusters are put into operation. This is made possible because the highspecific impulse thrusters deliver such low thrust that the drivethereof does not generate stresses that could damage fragile structuressuch as solar panels or solar concentrators, or deployable antennas andmasts.

In some cases, the antennas may be deployed only after the thrustershave stopped firing, if the thrusters are positioned so that they mightpollute or erode the active surfaces of the antennas. Conversely, bydeploying the solar panels, even before the thrusters begin to fire, itis possible for the high specific impulse thrusters to make effectiveuse of the electrical power produced by the solar panels.

BRIEF DESCRIPTION OF THE DRAWINGS

Other characteristics and advantages of the invention appear from thefollowing description of particular implementations of the invention,given as examples and made with reference to the accompanying drawings,in which:

FIG. 1 is a diagram showing a conventional way of reaching geostationaryorbit using an impulse at a single point;

FIG. 2 is a diagram showing a spiral method of reaching orbit with thethrusters operating continuously and starting from an initial orbit thatis circular;

FIG. 3 is a diagram showing how orbit can be reached by using thrustarcs, operating the thrusters discontinuously and starting from anelliptical initial orbit;

FIG. 4 is a perspective diagram showing a spiral type method of theinvention for reaching orbit with continuous thruster operation andstarting from an elliptical initial orbit;

FIGS. 5 and 6 are views of the FIG. 4 diagram respectively on a planeXOY and on a plane orthogonal to the XOY plane and containing an axis Z;

FIG. 7 is a diagrammatic perspective view of a first embodiment of asatellite to which the invention is applicable;

FIG. 8 is a diagrammatic perspective view of a second embodiment of asatellite to which the invention is applicable;

FIG. 9 is a diagram showing an elliptical orbit and defining the zone ofthrust around the half-ellipse centered on apogee;

FIG. 10 is a diagram showing an elliptical orbit and defining the zoneof thrust around the half-ellipse centered on perigee;

FIG. 11 is a diagram analogous to FIG. 5 showing a particular example ofthe method-of the invention as applied to reaching a geostationary orbitby a “similar” method;

FIG. 12 is a diagram analogous to the diagram of FIG. 4 showing aparticular example of the method of the invention as applied to reachinggeostationary orbit by an “alternative” method;

FIG. 13 is a block diagram of controlling and monitoring circuits in asystem of the invention;

FIG. 14 is a diagram applicable to a system of the invention showing howtotal thrust can be steered out from the orbital plane by modulating thethrust of one of the thrusters; and

FIG. 15 is a block diagram of a system of the invention using apropulsive system having high specific impulse thrusters of the closedelectron drift plasma type.

MORE DETAILED DESCRIPTION

Three conventional methods that have already been proposed for placing asatellite on a circular orbit, and they are recalled initially withreference to FIGS. 1 to 3.

FIG. 1 is a diagram of the various stages in the Hohmann maneuver,comprising a launch stage 1 during which the rocket places a satelliteon a low orbit, followed by a transfer stage 2 during which thesatellite is injected by an increase of speed at perigee onto anelliptical orbit whose apogee corresponds to the desired final altitude,and at which a second increase of speed 3 serves to inject the satelliteonto a final circular orbit 4. Portions referenced 5 and 6 representzones of intense radiation known as the Van Allen belts.

FIG. 2 shows the use of a spiral trajectory implementing accelerationthat is very low, but continuous. There can be seen the stage 1 oflaunching into low circular orbit, and the transfer stage 2 which isconstituted by the spiral orbit proper. Reference 7 shows the locationwhere spiral firing is stopped. Reference 3 shows the location of thefinal maneuver that serves to inject the satellite into the finalcircular orbit 4. The Van Allen belts are marked by references 5 and 6as in FIG. 1.

The drawbacks of the orbit changing methods shown in FIGS. 1 and 2 aredescribed above and are not described again.

FIG. 3 shows how an orbit is reached by applying a “thrust arc” method.

After the launch stage 1, the satellite is put on a first ellipticalhalf-orbit 2 and towards the apogee thereof thrusters are fired todeliver a speed increment 3A. The satellite then describes a secondelliptical orbit 4A having a perigee that is considerably higher and anapogee that is very slightly higher. Firing the thruster again addsanother speed increment 3B at the new apogee, thereby placing thevehicle on a new elliptical orbit 4B having a perigee that isconsiderably higher and an apogee that is slightly higher than in thepreceding orbit. This is continued until reaching elliptical orbit 4Nwhose apogee corresponds to the desired final altitude, enabling thevehicle to be placed on its final circular orbit 4.

As in FIGS. 1 and 2, the Van Allen belts are marked by references 5 and6.

The drawbacks of a method of reaching orbit by the technique ofsuccessive thrust arcs are already given above and are not describedagain.

With reference to FIGS. 4 to 6, there follows a description of a methodof the present invention for putting a space vehicle such as a satelliteinto orbit.

After a stage 101 of launching from the earth 100, by means of alauncher, a satellite is placed on an elliptical orbit 102.

Firing to put the satellite on its final orbit can be performed eitherimmediately, or else at the convenience of the user, after a certainnumber of revolutions on the initial orbit 102, e.g. when the satellitereaches a point 103.

At the moment when it is decided to start transferring the satellitesituated at point 103, a set of high specific impulse thrustersbelonging to the satellite is caused to fire, with the number of saidthrusters possibly being limited to one.

The high specific impulse thrusters are caused to operate continuously,and by steering their total thrust, they cause the satellite to describea spiral 104A, 104B, . . . , characterized by successive intermediateorbits in which apogee altitude increases more quickly than perigeealtitude.

Once the apogee altitude of an intermediate orbit is sufficient, e.g. atpoint 104G, the thrusters are kept in operation, and it is only therelationship controlling steering of total thrust that is modified inorder also to change inclination.

When the satellite describes portion 104K of the spiral trajectory, thethrusters are kept in operation and only the total thrust steeringrelationship is modified so that the spiral described then decreases inapogee altitude on each revolution, i.e. on each new intermediate orbit,while perigee altitude increases.

Once the desired change in inclination has been obtained, e.g. at point104N, the thrusters are kept in operation and only the total thruststeering relationship is modified to bring the thrust into alignmentwith the orbital plane while keeping it oriented in such a manner thatapogee altitude decreases while perigee altitude increases.

The spiral then continues by means of intermediate orbits 104N to 104P.The thrusters are caused to cease operating at point 107, for example,once the apogee and perigee altitudes are equal or approximately equalto the altitudes required by the target final orbit.

As can be seen from the example described above, it may be advantageousfor the inclination of the orbit not to be changed immediately onstarting the thruster(s). At the beginning of firing with continuousoperation of the thrusters, it is appropriate to allow perigee altitudeto increase quickly so as to minimize the time spent passing through theVan Allen belts. This time could be lengthened if inclination is to bechanged as well, since that would require a reduction in the thrustcomponent serving to increase perigee altitude.

Similarly, the desired change of inclination could be completed at point104N even before the target final orbit has been reached, so as tooptimize overall performance, given that the greater the apogeealtitude, the greater the effectiveness with which inclination ischanged.

The various stages of thruster operation during the spiral trajectorycan take into account the constraint of final positioning of the vehicleat its target point on the geostationary orbit, or in the vicinitythereof. Great latitude exists during the spiral trajectory which, atthe cost of a minor impact on performance, make it possible to reach anypoint on the geostationary orbit when the thrusters cease to operate.

It may be observed that the method of the invention has the particularadvantage of making it possible to define a maneuver during which thethrusters need to be started only once, and which suffices to go from aninitial orbit to a target final orbit. Naturally it is always possibleto interrupt firing of the thruster(s) temporarily, at the convenienceof the user, for reasons other than the needs of reaching the finalorbit, e.g. for performing maintenance operations on the trusters or thesatellite, or to satisfy orbit-mapping, telemetry, or remote controlrequirements.

Also, in certain applications, inclination can be changed as soon as thethrusters begin to-fire or right up to the end of thruster firing if,for example, it is desired to bring the satellite more quickly to itsfinal position on the geostationary orbit or to avoid passing throughthe geostationary orbit with zero inclination.

As is known, apogee and perigee are defined as being maximum and minimumdistances relative to the earth as reached by a space vehicle during onerevolution. In general, the orbit described by the satellite is close toelliptical in shape, even while the high specific impulse thrusters arein operation, given that their thrust is low. That is why it can beconsidered that the apogee and perigee in question can be approximatelythose of the ellipse osculating the mean orbital trajectory.

To obtain the looked-for effect during the stages of operationcorresponding to trajectory portions 104K to 104N, it is possible duringthe half-ellipse centered on apogee to steer thrust in the localhorizontal plane, as shown diagrammatically in FIG. 9, and to steerthrust during the half-ellipse centered on perigee in a plane orthogonalto the orbital plane and tangential to the orbit, in the oppositedirection to the speed, as shown diagrammatically in FIG. 10.

In FIGS. 9 and 10, there can be seen respectively the angle a of theapogee thrust arc and the angle α′ of the perigee thrust arc. In FIGS. 9and 10, the following parameters of the ellipse are marked:

a=semi major axis

b=semi minor axis

c=distance between the focus and the center of the ellipse.

The eccentricity of the ellipse is then given by the ratio c/a.

FIG. 11 shows how orbit can be reached using a “similar” method. Thismethod provides a very significant advantage concerning time spentpassing through the Van Allen belts. The initial orbit is an orbithaving a perigee of 620 kg and an apogee of 71,000 km. Continuous firingof the high specific impulse thrusters begins in the vicinity of thebeginning of the half-ellipse that is said to be “around” apogee.

The total thrust is directed in the local horizontal plane giving riseto increases both of perigee altitude and of apogee altitude. Then, onceperigee altitude exceeds 20,000 km (which value is a function, inparticular, of Van Allen belt activity, and the value taken intoconsideration could be as little as 8,000 km, for example) the directionof thrust is changed so as to be inertial (i.e. tangential to speed atapogee), thus having the effect of increasing perigee altitude and ofdecreasing apogee altitude; and then once the eccentricity of the orbitis zero, thrust direction is changed to be tangential and opposite tospeed, thus having the effect of reducing perigee altitude and ofreducing apogee altitude. As a result, apogee altitude and perigeealtitude are caused in the end to coincide with the altitude of thetarget orbit, i.e. the geostationary orbit. It may be observed that thenumber of orbits passing through the Van Allen belts is smaller than inthe case shown in FIG. 4.

FIG. 12 shows how orbit can be reached using the “alternative” method.This method provides a very significant advantage concerning time spentpassing through the Van Allen belts. The initial orbit has a perigee of620 km and an apogee of 330,000 km. Continuous firing of the highspecific impulse thrusters begins in the vicinity of the beginning ofthe halfellipse that is said to be “around” apogee. Total thrust isdirected to be inertial (tangential to speed at apogee), thus having theeffect of increasing perigee altitude and of decreasing apogee altitude.

Then, once the eccentricity of the orbit is zero, the thrust directionis changed so as to be in the local horizontal plane in the directionopposite to the speed, thereby reducing perigee altitude and reducingapogee altitude. As a result, perigee altitude and apogee altitudefinally coincide with the altitude of the target orbit, i.e. thegeostationary orbit. It may be observed that the number of orbitspassing through the Van Allen belt is much smaller than in the case ofFIG. 4, and is smaller than in the case of FIG. 11.

Reference is now made to FIGS. 7 and 8 which show two variants of asatellite 11 to which the invention is applicable. This space vehicle 11is fitted with high specific impulse thrusters 12 mounted on platformsor bases 13, at least some of which are steerable. The thrusters 12′ aredesigned to provide redundancy during the maneuver to reach orbit. Otherauxiliary thrusters 14 using the same matter as the thrusters 12 and 12′may also be mounted on the body of the satellite 11. The satellite isfitted with solar panels 10 each mounted on a structure 9 forming partof a solar panel drive mechanism referenced 9 a. The satellite is shownas having its axis OZ directed towards earth. The axis OZ is pointedtowards the center of the earth by infrared horizon centers 8 a ortowards a beacon by a radiofrequency detection 8 b. Auxiliary horizonsensors 15 may be implanted on other faces of the body of the satellite11.

The thrusters 12 and 12′ are also used for controlling orbit andattitude of the satellite during the operating lifetime of the vehicleas well as during the orbit-changing maneuver.

The system associated with the satellite for reaching its orbitcomprises a device for putting the thruster(s) into operation on acontinuous basis after the satellite has been placed on an initial orbitand until it has reached its target final orbit. The system alsoincludes a device for steering thrust by acting either on the steerablebases or else on other steering means available to the satellite (e.g.inertia wheels or kinetic wheels). Means are also provided forgenerating steering control such that starting from an ellipticalinitial orbit having a sidereal period that is different from and, forexample, less than that of the target final orbit:

at the beginning of the maneuver, on each successive revolution, theeffect of the thrusters 12 operating and the direction in which theyoperate cause the apogee of the orbit to increase together with a lesserincrease in the perigee of the orbit; and

at the end of the maneuver, on each successive revolution, the effect ofthe thrusters 12 operating is to decrease the apogee of the orbit and toincrease the perigee of the orbit.

The means provided for controlling steering can also be adapted toimplement the above-described “alternative” method, for example.

As shown diagrammatically in FIG. 13, various methods can be used tosteer the total thrust by the means that generate the steering controlsignals.

To steer thrust in a local horizontal plane, the steering frame ofreference may be given by an infrared horizontal sensor 15 a capable ofaccommodating varying distance between the earth and the satellite (e.g.a scanning sensor or a CCD matrix sensor). Third axis (yaw axis) controlmay be obtained in the general case by using a sun or star sensor 18 a,18 b, or 19. Aiming accuracy to within ±2° is more than sufficient forperforming the mission.

Thereafter, to steer thrust in a manner that is anti-tangential toorbital speed, the satellite rotates through about 180° about thenorth-south axis so as to reverse the thrust direction. It is thennecessary to have auxiliary horizon sensors 15 b and 15 c on twoopposite faces (east-west). The rotation is advantageously achieved bycreating torque by means of the kinetic wheel or the reaction wheel 17disposed on the north-south axis, and the amplitude of the rotation ismeasured by an integrating gyro 16, or by solar sensors 18 a and 18 b.

To steer the thrust in an inertial direction, a sun sensor 18 a, 18 b,or 18 c provide an attitude reference that moves substantially through1° per day relative to the stars. The onboard computer 121 computes theangle between the sun and the direction desired for total thrust andcontrols rotation of the satellite accordingly (by acting on the inertiawheel 17 via the attitude and orbit control system 122) until thedesired angle is reached.

If the sun sensor is situated on the solar panel 10 (sensors 18 a and 18b), the computer 121 causes the angle between the solar panels 10 andthe satellite body 11 to vary by acting on the solar panel drivemechanisms while the attitude and orbit control system 122 of thesatellite (AOCS) is solar controlled to point the panel towards the sun.

An attitude controlling torque can also be created by deliberatelypointing the steerable bases 13 supporting the high specific impulsethrusters 12 and 12′.

FIG. 13 shows a set of high specific impulse thrusters 12 started underthe control of a control circuit 120, a set of optional small auxiliarythrusters 14 using the same matter (e.g. xenon) as the set of thrusters12, a gyro 16, reaction wheels 17, a sun sensor 18 a for controlling thenorth solar panel drive mechanism, a sun sensor 18 b for controlling asouth solar panel drive mechanism, a set of sun sensors 19 secured tothe body of the satellite 11, plates 13 for supporting the high specificimpulse thrusters 12, a circuit 131 for controlling steering of theplates 13, an infrared horizon sensor 15 a, an onboard computer 121, andinterfaces and actuators 122 constituting a system for controllingattitude and orbit.

Optionally, and as also shown in FIG. 13, there may be an easthorizontal sensor 15 b and a west horizontal sensor 15 c.

To steer the total thrust away from the orbital plane (as is requiredwhen it is desired to change the inclination of successive orbits), itis possible to pivot the entire satellite about its yaw axis for example(axis OZ in FIG. 8).

If, as shown in FIGS. 7 and 8, a plurality of high specific thrustersare installed, e.g. two thrusters, it is also possible to steer totalthrust with a component lying outside the orbital plane by appropriatelysteering the steerable bases 13 and also modulating the intensity ofthrust from one of the two thrusters 12 so that in spite of thecomponent lying outside the orbital plane, the total thrust vectorpasses through the center of mass of the vehicle. This particularpossibility is shown diagrammatically in FIG. 14. To obtain total thrust(vector 23) passing through the center of mass 20 and having a componentlying outside the orbital plane, the steerable bases 13 are steereddifferently for the thrusters 12 on the north and on the south. Thrustfrom the north thruster 12 (vector 21) is adjusted so that the totalthrust of vector 23 which is equal to the vector sum of the vectors 21and 22 passes through the center of mass 20.

FIG. 15 is a block diagram of a propulsion system enabling orbit to bereached in application of the invention.

One or more tanks 31 of matter 30, e.g. xenon, feed(s) a manifold 32.The manifold 32 may contain a system for regulating pressure or flowrate together with a set of valves and sensors (not shown). Via nominallines 27 and redundant lines 37′, the manifold 32 delivers the matter 30to flow rate controllers 33 which feed the high specific impulsethrusters 12 and 12′. Optional auxiliary thrusters 14 using the-samematter 30 are fed from the lines 37 and 37′.

The thrusters 12 and 12′ are fed with electrical power via anelectricity distributor 35 that may contain a set of relays and thatcomes from an electricity conditioning module 36, itself directlyconnected to the electrical power source of the space vehicle-and to itscontrol and monitoring bus.

Each base 38 supporting two thrusters 12 and 12′ is steerable relativeto the space vehicle by command action on a respective steering device34.

The system of the invention is mainly applicable to transfers betweeninitial orbits that are elliptical such as GTO, and target final orbitsthat are circular such as GEO. However, it may also be advantageous,starting from an elliptical initial orbit GTO to reach some otherelliptical orbit that is inclined differently.

What is claim is:
 1. A system for placing a space vehicle, such as asatellite, on a target orbit such as that adapted to normal operation ofthe space vehicle and starting from an elliptical initial orbit thatdiffers substantially from the target orbit, and in particular that haseccentricity that is different from that of the target orbit, the systemcomprising: a set of platforms mounted on the space vehicle; a set ofthrusters mounted on said platforms and having high specific impulse,greater than 5000 Ns/kg, and low thrust, less than 10 N, to create atotal thrust force applied to the space vehicle; a control device forputting the thrusters into continuous operation after the space vehiclehas been placed on its initial orbit so as to enable said space vehicleto reach a target orbit via a completely spiral trajectory, ignoringpossible service interruptions, and to stop the thrusters firing whenthe apogee altitude, the perigee altitude, and the orbital inclinationof an intermediate orbit of the space vehicle are substantially equal tothe values of the target orbit; and a thrust direction control devicecomprising at least first means for operating, during a first stage ofcontinuous firing of the set of thrusters, to generate first controlsignals for aiming total thrust such that on each successive revolutionof the space vehicle and in each intermediate orbit, apogee altitudeincreases, perigee altitude increases to a lesser extent, and anydifference in the inclination of the intermediate orbit relative to thetarget orbit is decreased, and at least second means for operating,during a second stage of the set of thrusters firing, to generate secondcontrol signals for aiming total thrust such that on each successiverevolution of the space vehicle and in each intermediate orbit, apogeealtitude decreases, perigee altitude increases, and any difference ofinclination between the intermediate orbits and the target orbitdecreases.
 2. A system according to claim 1, wherein the set ofthrusters having high specific impulse comprises plasma type closedelectron drift thrusters, ion thrusters, or Arcjet type thrusters.
 3. Asystem according to claim 1, wherein said first and second means forgenerating the first and second control signals for aiming total thrustinclude a set of sensors enabling the attitude of the space vehicle tobe verified or checked.
 4. A system according to claim 1, wherein saidfirst means of said device for controlling thrust aiming include meansfor aligning total thrust in a local horizontal plane, mainly directedin the direction of the speed of the space vehicle.
 5. A systemaccording to claim 1, wherein said second means of said device forcontrolling thrust aiming comprise means for operating around apogee tobring the total thrust into alignment in a local horizontal plane mainlywith the speed direction of the space vehicle, and means for operatingaround perigee to bring the total thrust into alignment with a directionopposite to the orbital speed of the space vehicle in a plane that issubstantially orthogonal to the orbital plane.
 6. A system according toclaim 5, wherein said means for operating around apogee to align totalthrust serve to aim total thrust substantially along the osculatinghalf-ellipse centered on apogee.
 7. A system according to claim 5,wherein said means for operating around perigee to align total thrustserve to aim total thrust substantially on the osculating half-ellipsecentered on perigee.
 8. A system according to claim 1, wherein saidfirst and second means of said device for controlling thrust aiminginclude means for aligning thrust in a plane passing via directions thatare fixed or almost fixed in space and passing almost perpendicularly tothe plane that is tangential to the orbit.
 9. A system according toclaims 1, including means for aiming total thrust that are constitutedby means belonging to the space vehicle such as inertia wheels of asystem for controlling the attitude of the space vehicle.
 10. A systemaccording to claim 1, including means for aiming total thrustconstituted by at least some of said platforms for supporting sets ofthrusters, said platforms being steerable.
 11. A system according toclaim 10, wherein said steerable platforms can be steered through morethan 10° at least about one axis.
 12. A system according to claim 10,including means for steering total thrust including means fordifferentially aiming said steerable platforms and means forservo-controlling the thrust from each thruster to a predeterminedvalue, thereby enabling total thrust to be generated that passes throughthe center of mass of the space vehicle and that has a component lyingoutside the orbital plane.
 13. A system according to claim 1, whereinsaid thrusters having high specific impulse also constitute means forcontrolling attitude and orbit of the space vehicle such as a satellite.